This invention relates to a new and improved thruster and, more specifically, to an electrothermal thruster adapted for maneuvering spacecraft.
Electrically heated hydrazine thrusters are well known, and are discussed in:
AIAA 8th Electrical Propulsion Conference 1970 #70-1161;
AIAA/SAE 8th Joint Propulsion Specialist Conference 1972 #72-1152;
AIAA 9th Electricl Propulsion Conference 1972 #72-451; U.S. Pat. No. 3,081,595;
"Study of Monopropellants for Electrothermal Hydrazine Thrusters" March 1973-May 1974, by J. D. Kuenzly, Final Report for Goddard Space Flight Center, Contract NAS 5-23202; and
International Conference--Properties of Hydrazine and its Potential Applications as an Energy Source (Poitiers, France, Oct. 21-25, 1974)-"Electrothermal Hydrazine Thruster Development for Low Thrust Applications."
These thrusters have particular capability to provide one or more of the following functions: attitude control, initial orbit correction, initial stationkeeping, repositioning, drag make-up, orbit raising and evasive maneuvers.
As an example, a commercial communications satellite would require all of the above functions (except the last three) to be performed by the on-board propulsion system. For such a spacecraft with an orbital weight of 2000 pounds, an on-board catalytic hydrazine system would weigh about 500 pounds for a seven-year mission. The total payload weight would be about 400 pounds with the remainder of the weight allocated to the power system, structures, command and telemetry and so forth.
In addition to satellites for commercial communications, spacecraft for military, or scientific purposes are typical of long life earth orbital spacecraft which require thie maneuvering capability.
Obviously, any reduction in propulsion system weight which could be used to increase the payload capability would be highly desirable. Increasing the payload would permit a greater return on the spacecraft program investment through such means as greater communications capacity, the accumulation of more scientific data, the provision of additional surveillance equipment, fewer launches per series, or the capability to use smaller, less expensive launch vehicles.
Initially, in the United States space program, virtually all spacecraft which required on-board propulsion relief on cold gas systems which produced a specific impulse of about 70 seconds. However, when spacecraft missions became longer and more complex, the demands placed on spacecraft propulsion systems become more stringent. As the propulsion requirements increased, cold gas was supplanted by catalytically-decomposed hydrogen peroxide which offered significant propulsion system weight savings at the cost of increased complexity and various operational difficulties. The use of hydrogen peroxide propulsion systems was discontinued in favor of catalytically-decomposed anhydrous hydrazine which offered still greater weight savings and eliminated some of the hydrogen peroxide operation problems. Catalytic hydrazine thrusters typically deliver a specific impulse of 215-235 seconds (steady state) and systems employing such thrusters are now widely used for a variety of spacecraft missions.
Specific impulse is a figure-of-merit commonly used for thruster performance and is defined as the thrust that can be obtained with a propellant weight flow rate of unity. Specific impulse is determined by the energy content of the gases in the thrust chamber and the efficiency of the nozzle expansion process which converts this chamber energy to kinetic energy in the exhaust. The relationship between chamber energy and specific impulse is ##EQU1## where H is that portion of the energy (enthalpy) which is converted to kinetic energy.
In a catalytic hydrazine thruster, the decomposition of anhydrous hydrazine into hydrogen and nitrogen is assumed to occur in two stages, viz., an initial decomposition and final dissociation, as follows: EQU 3N.sub.2 H.sub.4 .fwdarw.4NH.sub.3 +N.sub.2 +144,300 BTU (exothermic decomposition)
and EQU 4NH.sub.3 .fwdarw.2N.sub.2 +6H.sub.2 -79,200 BTU (endothermic dissociation)
The result of these reactions produces a mixture of N.sub.2, H.sub.2, and NH.sub.3 at a temperature of about 1600.degree.-1800.degree. F. Isentropic expansion of these decomposition products through a nozzle results in a theoretical specific impulse of 210 to 260 seconds. However, when various performance loss mechanisms (such as heat losses, incomplete expansion, nozzle divergence and so forth) are considered, the delivered steady state specific impulse of current hydrazine thrusters is usually in the range of 215 to 235 seconds. Consequently, the potential for performance improvement of catalytic hydrazine thrusters is very limited since the chamber enthalpy (and theoretical impulse) is limited by the net chemical energy released by the decomposition of the propellant.
Hence, there is required an increase in thruster capability to at least 300 Isp while still maintaining a sensibly constant delivered power with no significant weight increase of either thruster or propellant.